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HQ 3.0/8 AIRFOIL (hq308-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: HQ 3.0/8 AIRFOIL (hq308-il)
Reynolds number: 100,000
Max Cl/Cd: 62.54 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hq308-il-100000.txt
Download as CSV file: xf-hq308-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HQ 3.0/8 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3856   0.09511   0.09046  -0.0340   1.0000   0.0672
  -7.750  -0.3962   0.09334   0.08881  -0.0347   1.0000   0.0676
  -7.500  -0.4062   0.09113   0.08672  -0.0386   1.0000   0.0679
  -7.250  -0.4120   0.08828   0.08390  -0.0426   1.0000   0.0681
  -7.000  -0.3958   0.08367   0.07931  -0.0306   1.0000   0.0733
  -6.750  -0.3999   0.08116   0.07687  -0.0305   1.0000   0.0756
  -6.500  -0.4051   0.07844   0.07422  -0.0317   1.0000   0.0778
  -6.250  -0.4097   0.07534   0.07111  -0.0370   1.0000   0.0805
  -6.000  -0.4113   0.07262   0.06817  -0.0445   1.0000   0.0820
  -5.750  -0.4086   0.06816   0.06395  -0.0379   1.0000   0.0843
  -5.500  -0.4044   0.06551   0.06134  -0.0362   1.0000   0.0878
  -5.250  -0.3902   0.06192   0.05728  -0.0447   1.0000   0.0959
  -5.000  -0.3860   0.05804   0.05367  -0.0410   1.0000   0.0986
  -4.750  -0.3674   0.05498   0.05020  -0.0455   1.0000   0.1098
  -4.500  -0.3597   0.05157   0.04701  -0.0429   1.0000   0.1142
  -4.250  -0.3423   0.04826   0.04354  -0.0447   1.0000   0.1267
  -4.000  -0.3246   0.04531   0.04048  -0.0456   1.0000   0.1412
  -3.750  -0.3041   0.04262   0.03757  -0.0473   1.0000   0.1657
  -3.500  -0.2872   0.04013   0.03511  -0.0469   1.0000   0.1857
  -3.000  -0.1886   0.03134   0.02389  -0.0520   1.0000   0.0685
  -2.750  -0.1570   0.02758   0.01963  -0.0523   1.0000   0.0557
  -2.500  -0.1302   0.02558   0.01735  -0.0521   1.0000   0.0535
  -2.250  -0.1036   0.02386   0.01522  -0.0520   1.0000   0.0564
  -2.000  -0.0782   0.02238   0.01363  -0.0517   1.0000   0.0595
  -1.750  -0.0527   0.02118   0.01227  -0.0511   1.0000   0.0620
  -1.500  -0.0269   0.02028   0.01122  -0.0505   0.9998   0.0678
  -1.250   0.0124   0.01893   0.01013  -0.0528   0.9955   0.1022
  -1.000   0.0335   0.01572   0.00992  -0.0498   0.9938   0.8439
  -0.750   0.0786   0.01574   0.00958  -0.0535   0.9849   1.0000
  -0.500   0.1201   0.01622   0.00969  -0.0569   0.9771   1.0000
  -0.250   0.1646   0.01672   0.00990  -0.0608   0.9697   1.0000
   0.000   0.2023   0.01708   0.01004  -0.0634   0.9602   1.0000
   0.250   0.2436   0.01748   0.01023  -0.0665   0.9517   1.0000
   0.500   0.2861   0.01783   0.01044  -0.0698   0.9431   1.0000
   0.750   0.3217   0.01812   0.01063  -0.0718   0.9329   1.0000
   1.000   0.3607   0.01840   0.01082  -0.0742   0.9236   1.0000
   1.250   0.4083   0.01855   0.01092  -0.0781   0.9152   1.0000
   1.500   0.4469   0.01862   0.01097  -0.0801   0.9033   1.0000
   1.750   0.4868   0.01862   0.01097  -0.0822   0.8915   1.0000
   2.000   0.5274   0.01857   0.01097  -0.0844   0.8804   1.0000
   2.250   0.5726   0.01838   0.01083  -0.0872   0.8707   1.0000
   2.500   0.6190   0.01799   0.01051  -0.0898   0.8615   1.0000
   2.750   0.6568   0.01761   0.01021  -0.0907   0.8484   1.0000
   3.000   0.6937   0.01714   0.00988  -0.0911   0.8347   1.0000
   3.250   0.7284   0.01669   0.00951  -0.0911   0.8205   1.0000
   3.500   0.7555   0.01647   0.00939  -0.0901   0.8024   1.0000
   3.750   0.7859   0.01610   0.00912  -0.0893   0.7848   1.0000
   4.000   0.8176   0.01563   0.00881  -0.0886   0.7666   1.0000
   4.250   0.8436   0.01538   0.00866  -0.0871   0.7421   1.0000
   4.500   0.8702   0.01511   0.00846  -0.0856   0.7144   1.0000
   4.750   0.8951   0.01497   0.00837  -0.0839   0.6801   1.0000
   5.000   0.9193   0.01492   0.00833  -0.0821   0.6376   1.0000
   5.250   0.9418   0.01506   0.00845  -0.0802   0.5842   1.0000
   5.500   0.9620   0.01549   0.00864  -0.0780   0.5195   1.0000
   5.750   0.9798   0.01626   0.00906  -0.0758   0.4500   1.0000
   6.000   0.9965   0.01723   0.00968  -0.0737   0.3867   1.0000
   6.250   1.0135   0.01823   0.01042  -0.0719   0.3315   1.0000
   6.500   1.0307   0.01925   0.01127  -0.0702   0.2830   1.0000
   6.750   1.0473   0.02039   0.01223  -0.0685   0.2388   1.0000
   7.000   1.0620   0.02167   0.01332  -0.0666   0.1868   1.0000
   7.250   1.0744   0.02327   0.01468  -0.0646   0.1330   1.0000
   7.500   1.0884   0.02479   0.01597  -0.0629   0.0883   1.0000
   7.750   1.0988   0.02706   0.01806  -0.0604   0.0578   1.0000
   8.000   1.1096   0.03035   0.02122  -0.0578   0.0435   1.0000
   8.250   1.1311   0.03298   0.02408  -0.0562   0.0374   1.0000
   8.500   1.1528   0.03679   0.02791  -0.0556   0.0331   1.0000
   8.750   1.1709   0.03945   0.03102  -0.0540   0.0306   1.0000
   9.000   1.1862   0.04234   0.03435  -0.0522   0.0285   1.0000
   9.250   1.1974   0.04608   0.03857  -0.0501   0.0279   1.0000
   9.500   1.2025   0.05012   0.04312  -0.0476   0.0281   1.0000
   9.750   1.2013   0.05430   0.04779  -0.0448   0.0284   1.0000
  10.000   1.1950   0.05848   0.05238  -0.0419   0.0288   1.0000
  10.250   1.1826   0.06235   0.05659  -0.0387   0.0292   1.0000
  10.500   1.1663   0.06611   0.06062  -0.0359   0.0296   1.0000
  10.750   1.1482   0.07012   0.06488  -0.0340   0.0299   1.0000
  11.000   1.1289   0.07456   0.06953  -0.0333   0.0302   1.0000
  11.250   1.1086   0.07951   0.07468  -0.0338   0.0304   1.0000
  11.500   1.0881   0.08508   0.08042  -0.0356   0.0307   1.0000
  11.750   1.0669   0.09150   0.08700  -0.0389   0.0309   1.0000
  12.000   1.0460   0.09883   0.09445  -0.0434   0.0312   1.0000
  12.250   1.0266   0.10705   0.10275  -0.0490   0.0316   1.0000
  12.500   1.0096   0.11589   0.11165  -0.0547   0.0320   1.0000
  12.750   0.8345   0.13121   0.12746  -0.0597   0.0423   1.0000
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